I. Field of the Invention
The present invention relates generally to turbine inlet nozzles for gas turbine engines and, more particularly, to such a turbine inlet nozzle with novel cooling means.
II. Description of the Prior Art
The previously known gas turbine engines typically comprise a support housing in which a main shaft is rotatably mounted. A compressor means is operatively secured to the shaft which receives air from an air inlet, compresses the air and supplies its compressed air outlet to a combuster assembly within the support housing. In the combustion assembly, fuel is mixed with the compressed air and ignited and the resulting combustion products from the combustor assembly expand through one or more turbine stages secured to the main shaft to rotatably drive the compressor and also to provide the engine output.
In the previously known designs for turbine engines, a nozzle is disposed in the gas stream passageway between the outlet from the combustor assembly and the first turbine stage. Such nozzles typically comprise a plurality of circumferentially spaced vanes positioned coaxially around the turbine shaft.
The trend in turbine engine design in recent years has been to increase the combustion temperature of the engine since the increase of combustion temperature likewise increases the overall efficiency of the turbine engine. Because of this, the turbine nozzle assemblies are subjected to extremely high temperatures. The materials conventionally used in the construction of the turbine inlet nozzle are incapable of withstanding the high temperature environment from the engine gas stream without excessive and unacceptable thermal distortion and must therefore be cooled during the operation of the turbine engine.
One previously known method of cooling the turbine nozzle assembly during the operation of the turbine engine has been to provide an impingement plate coaxially around the outer shroud of the nozzle assembly. The impingement plate includes a plurality of holes formed through it and is spaced radially outwardly from the outer perhiphery of the outer shroud thus forming a chamber therebetween. A portion of the compressed air outlet from the turbine engine compressor is diverted to the outer side of the impingement plate so that the compressed air flows through the holes in the impingement plate and against the outer shroud thus cooling it. Typically this cooling air flow is subsequently exhausted to the turbine engine gas stream and expelled from the turbine engine. In addition, a number of previously known turbine engines included hollow vanes through which the cooling air flow passed prior to its exhaustion into the turbine engine gas stream.
The use of impingement plates for cooling the turbine nozzle, however, have suffered from a number of disadvantages. One such disadvantage is the impingement plate is normally maintained at a much cooler temperature than the turbine nozzle components due to the cooling air flow through the impingement plate and thus the impingement undergoes differential thermal growth with respect to the other turbine nozzle components. Consequently, it has not been previously possible to secure the impingement plate directly to the shroud due to the high thermal stress and possible rupture between the impingement plate and the other nozzle components.
In order to prevent coolant leakage between the impingement plate and the nozzle components, it has been the previous practice to employ seals between the impingement plate and the nozzle shroud. In practice, however, such seals have proven inadequate thus resulting in high coolant loss, insufficient cooling of the turbine nozzle and degradation in the performance of the turbine engine. When the turbine nozzle is of multi piece construction, i.e. the shrouds and vanes are separately constructed and thereafter secured together, seals are also employed between the turbine nozzle components and these seals likewise result in high coolant leakage.
A still further disadvantage of the previously known turbine nozzles with impingement plate cooling means is the holes through the impingement plate are evenly distributed across it. Such impingement plates thus provide even cooling across the turbine engine nozzle and particularly the outer shroud. Advanced turbine engines, however, operate in an environment with varying gas temperature distribution and surface heating rate of the nozzle components. The varying gas distribution is due primarily to the characteristics of the combustion assembly while variation of the surface heating rate of the nozzle components is controlled primarily by the pressure distribution and flow shape across the nozzle surfaces. The previously known nozzle cooling means utilizing impingement plates have failed to address the varying cooling requirements of the turbine inlet nozzle and thus have failed to evenly cool the nozzle assembly in the desired fashion.